The present invention relates generally to a gas turbine engine component and more particularly to a nozzle segment having an integral outer band and shroud segment.
Gas turbine engines have a stator and one or more rotors rotatably mounted on the stator. The engines generally include a high pressure compressor for compressing flowpath air traveling through the engine, a combustor downstream from the compressor for heating the compressed air, and a high pressure turbine downstream from the combustor for driving the high pressure compressor. Further, the engines include a low pressure turbine downstream from the high pressure turbine for driving a fan positioned upstream from the high pressure compressor.
Downstream from the combustor, flowpath air temperatures are hot resulting in the components forming the flowpath being hot. As components reach these elevated flowpath air temperatures, their material properties decrease. To combat this reduction in material properties, flowpath air is extracted from cooler areas of the engine such as the compressor and blown through and around the hotter components to lower their temperatures. Delivering cooling air to the hotter components increases their lives, but extracting flowpath air from the cooler areas of the engine reduces the efficiency of the engine. Thus, it is desirable to minimize the amount of cooling air required by the hotter components to increase overall engine efficiency. In particular, it is important to minimize the cooling air introduced downstream from the nozzle throat. Cooling air introduced downstream from the nozzle throat is significantly more detrimental to engine performance than air introduced upstream from the nozzle throat.
FIG. 1 illustrates a conventional high pressure turbine nozzle assembly, designated in its entirety by the reference character 10. The nozzle assembly 10 includes nozzle segments, generally designated by 12, mounted on a nozzle support 14. Shroud segments 16 are mounted on a shroud hanger 18 downstream from the nozzle segments 12. The shroud hanger 18 is mounted on a support 20 surrounding the hanger. The nozzle segments 12 include an outer band segment 22 extending circumferentially around a centerline 24 of the engine having an inner surface 26 forming a portion of an outer flowpath boundary. A plurality of nozzle vanes 28 extend inward from the outer band segment 22 and an inner band segment 30 extends circumferentially around the inner ends of the nozzle vanes. The inner band segment 30 has an outer surface 32 forming a portion of an inner flowpath boundary of the engine. A rotating disk 34 and blades 36 are mounted downstream from the nozzle segments 12 inside the shroud segments 16.
Cooling air is introduced into two cavities 38, 40 positioned outboard from the nozzle outer band segments 22 and the shroud hanger 18, respectively. Part of the cooling air delivered to the cavity 38 outboard from the outer band segments 22 enters passages 42 in the nozzle vanes 28 and exits through cooling holes 44 formed in the surface of the vanes to cool the vanes by film cooling. Some of the cooling air delivered to the cavity 38 leaks into the flowpath between the circumferential ends of the outer band segments 22 and some of the cooling air leaks into the flowpath past a seal 46 positioned between the nozzle outer band segments and the shroud hanger 18. The cooling air delivered to the cavity 40 positioned outboard from the shroud hangers 18 impinges upon the shroud segments 16 to cool them by impingement cooling and then leaks into the flowpath between the circumferential ends of the shroud segments.
Among the several features of the present invention may be noted the provision of a gas turbine engine component. The component comprises a nozzle outer band extending circumferentially around a centerline of the engine having an inner surface forming a portion of an outer flowpath boundary of the engine. Further, the component includes a plurality of nozzle vanes extending inward from the outer band. Each of the vanes extends generally inward from an outer end mounted on the outer band to an inner end opposite the outer end. In addition, the component comprises an inner band extending circumferentially around the inner ends of the plurality of nozzle vanes having an outer surface forming a portion of an inner flowpath boundary of the engine. Still further, the component includes a shroud integral with the outer band extending circumferentially around the centerline of the engine and having an inner surface forming a portion of the outer flowpath boundary of the engine adapted for surrounding a plurality of blades mounted in the engine for rotation about the centerline thereof.
In another aspect, the present invention includes a high pressure turbine nozzle segment for use in a gas turbine engine. The nozzle segment comprises an outer band segment extending circumferentially around a centerline of the nozzle segment and rearward to a shroud segment integrally formed with the outer band segment extending circumferentially around the centerline. The outer band segment and shroud segment have a substantially continuous and uninterrupted inner surface forming a portion of the outer flowpath boundary of the engine. The nozzle segment also includes nozzle vanes extending inward from the outer band segment. Each of the vanes extends generally radially inward from an outer end mounted on the outer band segment to an inner end opposite the outer end. In addition, the nozzle segment comprises an inner band segment extending circumferentially around the inner ends of the nozzle vanes having an outer surface forming a portion of an inner flowpath boundary of the engine.
Other features of the present invention will be in part apparent and in part pointed out hereinafter.